Present embodiments relate generally to rotor blades for gas turbine engines. More particularly, but not by way of limitation, present embodiments relate to ceramic matrix composite ply architecture for platform and damper retaining features integrally formed with the rotor blades.
In the gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine includes a first stage nozzle and a rotor assembly having a disk and a plurality of turbine blades. The high pressure turbine first receives the hot combustion gases from the combustor and includes a first stage stator nozzle that directs the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a first rotor disk. In a multi-stage turbine, a second stage stator nozzle is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second rotor disk. The stator nozzles direct the hot combustion gas in a manner to maximize extraction at the adjacent downstream turbine blades.
The high pressure rotor discs are joined to the compressor rotors by a corresponding high pressure shaft for powering the compressor during operation. A multi-stage low pressure turbine follows the multi-stage high pressure turbine and is typically joined by a low pressure shaft to low pressure compressor and a fan disposed upstream from the low pressure compressor in a typical turbofan aircraft engine configuration.
As the combustion gases flow downstream through the turbine stages, energy is extracted therefrom and the pressure of the combustion gas is reduced. The combustion gas is used to power the compressor as well as a turbine output shaft for power. In this manner, fuel energy is converted to mechanical energy of the rotating shaft to power the compressor and supply compressed air needed to continue the process.
Turbine rotor assemblies typically include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that having a pressure side and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade may also include a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. Known blades are hollow such that an internal cooling cavity is defined at least partially by the airfoil, platform, shank, and dovetail.
These prior art rotor blades have been formed of metallic materials. However increased performance goals have resulted in a goal of using alternate materials to allow for higher turbine inlet temperatures. Improved temperature capability has been accomplished through the use of ceramic matrix composites in the gas turbine engine components. However, the integration of metallic or other material components with the ceramic matrix composite components has been problematic. For example, with respect to rotor blades, it is desirable to retain sheet metal dampers in determinant, robust manner during all operating conditions as well as making any damper retaining features integral with the CMC rotor blades.
As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies in order to allow blade assembly of CMC with integral platform and damper retaining features.